An airframe traditionally has a metal, mainly aluminium alloy, fuselage put together by assembling frames, stringers and panels. Areas for probes, doors, other openings and the like are cut out from the panels of the fuselage. The skin of the fuselage is then generally reinforced internally, around the cut-out, by a stepped local increase in thickness.
If these thick (conventionally greater than 4 mm thick) reinforcing areas suffer serious damage—following an incident in service, fire, corrosion, etc.—the damaged skin is conventionally cut away and a thick external plate affixed.
The use of external plates is described for example in U.S. Pat. No. 4,517,038. In that document it is recommended to cut away the damaged part to form a preselected opening, to introduce into this opening an assembly—composed of a stack of layers of fibres, of substantially the same size as the opening, joined to a metal plate of greater size—and to fix the plate by means of rivets to the skin of the damaged structure. Such a repair is not reliable, in particular with regard to fire resistance, and causes aerodynamic problems, interfering with probes if it is located near a probe.
It is known from the patent document EP 0 471 923 to repair a damaged non-load-bearing airframe structure by fixing an external fire-resistant plate constituted of a layer of woven glassfibre fabric impregnated with a thermosetting resin. The plate is fixed by a plurality of mechanical clamping means.
If environmental constraints multiply, this type of repair is difficult to apply: for aerodynamic reasons the thickness of the external plates must be limited or even eliminated in the vicinity of probes so as not to interfere with the flow of air, as such interference may falsify the measurements from the probes.
The plate is shaped manually or machined from thick sheet metal. Reworking is generally necessary to reduce the thickness of its contour along the lines of external fixing of the plate. This method is therefore lengthy and difficult and induces internal mechanical stresses linked to the fabrication process.
Moreover, during assembly, there are additional installation stresses, linked mainly to manufacturing tolerances. Moreover, misalignment of the neutral fibres also induces secondary bending stresses at the junction.
Finally, the repair may be limited in time with additional intermediate inspections: the final solution is always to replace the entire panel, which is the worst case scenario in terms of cost and down time.